This invention relates generally to turbomachines, such as turbine engines. More specifically, the invention is directed to methods and articles for impeding the flow of gas (e.g., hot gas) through selected regions of stator-rotor assemblies in turbomachines.
The typical design of most turbine engines is well-known in the art. They include a compressor for compressing air that is mixed with fuel. The fuel-air mixture is ignited in an attached combustor, to generate combustion gases. The hot, pressurized gases, which in modern engines can be in the range of about 1100 to 2000° C., are allowed to expand through a turbine nozzle, which directs the flow to turn an attached, high-pressure turbine. The turbine is usually coupled with a rotor shaft, to drive the compressor. The core gases then exit the high pressure turbine, providing energy downstream. The energy is in the form of additional rotational energy extracted by attached, lower pressure turbine stages, and/or in the form of thrust through an exhaust nozzle.
More specifically, thermal energy produced within the combustor is converted into mechanical energy within the turbine, by impinging the hot combustion gases onto one or more bladed rotor assemblies. (Those versed in the art understand that the term “blades” is usually part of the lexicon for aviation turbines, while the term “buckets” is typically used when describing the same type of component for land-based turbines). The rotor assembly usually includes at least one row of circumferentially-spaced rotor blades. Each rotor blade includes an airfoil that includes a pressure side and a suction side. Each airfoil extends radially outward from a rotor blade platform. Each rotor blade also includes a dovetail that extends radially inward from a shank extending between the platform and the dovetail. The dovetail is used to mount the rotor blade within the rotor assembly to a rotor disk or spool.
As known in the art, the rotor assembly can actually be considered as a portion of a stator-rotor assembly. The rows of rotor blades on the rotor assembly and the rows of stator vanes on the stator assembly extend alternately across an axially oriented flowpath for “working” the combustion gases. The jets of hot combustion gas leaving the vanes of the stator element act upon the turbine blades, and cause the turbine wheel to rotate in a speed range of about 3000-15,000 rpm, depending on the type of engine. (Again, in terms of parallel terminology, the stator element, i.e., the element which remains stationary while the turbine rotates at high speed, can also be referred to in the art as the “nozzle assembly”).
As depicted in the figures described below, the opening at the interface between the stator element and the blades or buckets can allow hot core gas to exit the hot gas path and enter the wheel-space of the turbine engine. In order to limit this leakage of hot gas, the blade structure typically includes axially projecting angel wing seals. According to a typical design, the angel wings cooperate with projecting segments or “discouragers” which extend from the adjacent stator element, i.e., the nozzle. The angel wings and the discouragers overlap (or nearly overlap), but do not touch each other, thus restricting gas flow. The effectiveness of the labyrinth seal formed by these cooperating features is critical for limiting the ingestion of hot gas into undesirable sections of the engine. The angel wings can be of various shapes, and can include other features, such as radial teeth. Moreover, some engine designs use multiple, overlapping angel wing-discourager seals.
A gap remains at the interface between adjacent regions of the nozzle and turbine blade, e.g., between the adjacent angel wing-discourager projections, when such a seal is used. The presence of the gap is understandable, i.e., the clearance necessary at the junction of stationary and rotating components. However, the gap still provides a path which can allow hot core gas to exit the hot gas path into the wheel-space area of the turbine engine.
As alluded to above, the leakage of the hot gas by this pathway is disadvantageous for a number of reasons. First, the loss of hot gas from the working gas stream causes a resultant loss in energy available from the turbine engine. Second, ingestion of the hot gas into turbine wheel-spaces and other cavities can damage components which are not designed for extended exposure to such temperatures, such as the nozzle structure support and the rotor wheel.
One well-known technique to further minimize the leakage of hot gas from the working gas stream involves the use of coolant air, i.e., “purge air”, as described in U.S. Pat. No. 5,224,822 (Lenehan et al). In a typical design, the air can be diverted or “bled” from the compressor, and used as high-pressure cooling air for the turbine cooling circuit. Thus, the coolant air is part of a secondary flow circuit which can be directed generally through the wheel-space cavity and other inboard regions. In one specific example, the coolant air can be vented to the rotor/stator interface.
Thus, the coolant air can function to maintain the temperature of certain engine components under an acceptable limit. However, the coolant air can serve an additional, specific function when it is directed from the wheel-space region into one of the gaps described previously. This counter-flow of coolant air into the gap provides an additional barrier to the undesirable flow of hot gas out of the gap and into the wheel-space region.
While coolant air from the secondary flow circuit is very beneficial for the reasons discussed above, there are drawbacks associated with its use as well. For example, the extraction of air from the compressor for high pressure cooling and cavity purge air consumes work from the turbine, and can be quite costly in terms of engine performance. Moreover, in some engine configurations, the compressor system may fail to provide purge air at a sufficient pressure during at least some engine power settings. Thus, hot gases may still be ingested into the wheel-space cavity.
It should be apparent from this discussion that new techniques for reducing the leakage of hot gases from a hot gas flow path into undesirable regions within a turbine engine or other type of turbomachine would be welcome in the art. Moreover, reduction of the cooling and cavity purge-air flow which is typically required to reduce the hot gas leakage would itself have other important benefits. For example, higher core air flow would be possible, thereby increasing the energy available in the hot gas flow path.
New techniques for accomplishing these goals must still adhere to the primary design requirements for a gas turbine engine or other type of turbomachine. In general, overall engine efficiency and integrity must be maintained. Any change made to the engine or specific features within the engine must not disturb or adversely affect the overall hot gas and coolant air flow fields. Moreover, the contemplated improvements should not involve manufacturing steps or changes in those steps which are time-consuming and uneconomical. Furthermore, the improvements should be adaptable to varying'designs in engine construction, e.g., different types of stator-rotor assemblies. It would also be very advantageous if the improvements were adaptable to the containment of lower-temperature gases (e.g., room temperature), as well as hot gases.